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  • 1
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    Boundary-layer stability analysis of Langley Research Center 8-foot LFC experimental data (1987)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: An analytical study of linear-amplifying instabilities of a laminar boundary layer as found in the experimental data of the LaRC/8-foot laminar-flow control (LFC) experiment was completed and the results are presented. The LFC airfoil used for this experiment was a swept, supercritical design which removed suction air through spanwise slots. The amplification of small disturbances by linear processes on a swept surface such as this can be due to either Tollmien-Schlichting (TS) and/or crossflow (CF) mechanisms. This study consists of the examination of these two instabilities by both the commonly used incompressible (SALLY and MARIA) analysis and the more involved compressible (COSAL) analysis. A wide range of experimental test conditions with variations in Mach number, Reynolds number, and suction distributions were available for this study. Experimentally determined transition locations were found from thin-film techniques and were used to correlate the n-factors at transition for the range of test cases.
    Keywords: AERODYNAMICS
    Type: Research in Natural Laminar Flow and Laminar-Flow Control, Part 2; p 471-489
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Helicopter blade dynamic loads measured during performance testing of two scaled rotors (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A test to determine the performance differences between the 27-percent-scale models of two rotors for the U.S. Army AH-64 helicopter was conducted in the Langley 14- by 22-Foot Subsonic Tunnel. One rotor, referred to as the baseline rotor, simulated the geometry and dynamic characteristics of the production baseline rotor, and the other rotor, referred to as the advanced rotor, was designed to have improved hover performance. During the performance test, the dynamic pitch-link forces and blade bending and torsion moments were also measured. Dynamic data from the forward flight investigation are reduced and presented. The advanced blade set was designed to have dynamic characteristics similar to those of the baseline rotor so that test conditions would not be limited by potential rotor instability and blade resonances, and so that the measured performance increments could be considered to be due purely to aerodynamic causes. Data show consistent trends with advance ratio for both blade sets with generally higher oscillatory loads occurring for the advanced blade set when compared with the baseline blade set.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89053 , L-16245 , NAS 1.15:89053 , AVSCOM-TM-87-B-7 , AD-A182870
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Description of the US Army small-scale 2-meter rotor test system (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A small-scale powered rotor model was designed for use as a research tool in the exploratory testing of rotors and helicopter models. The model, which consists of a 29 hp rotor drive system, a four-blade fully articulated rotor, and a fuselage, was designed to be simple to operate and maintain in wind tunnels of moderate size and complexity. Two six-component strain-gauge balances are used to provide independent measurement of the rotor and fuselage aerodynamic loads. Commercially available standardized hardware and equipment were used to the maximum extent possible, and specialized parts were designed so that they could be fabricated by normal methods without using highly specialized tooling. The model was used in a hover test of three rotors having different planforms and in a forward flight investigation of a 21-percent-scale model of a U.S. Army scout helicopter equipped with a mast-mounted sight.
    Keywords: AERODYNAMICS
    Type: NASA-TM-87762 , L-16165 , NAS 1.15:87762 , AVSCOM-TM-86-B-4 , AD-A178047
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Incompressible boundary-layer stability analysis of LFC experimental data for sub-critical Mach numbers (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An incompressible boundary-layer stability analysis of Laminar Flow Control (LFC) experimental data was completed and the results are presented. This analysis was undertaken for three reasons: to study laminar boundary-layer stability on a modern swept LFC airfoil; to calculate incompressible design limits of linear stability theory as applied to a modern airfoil at high subsonic speeds; and to verify the use of linear stability theory as a design tool. The experimental data were taken from the slotted LFC experiment recently completed in the NASA Langley 8-Foot Transonic Pressure Tunnel. Linear stability theory was applied and the results were compared with transition data to arrive at correlated n-factors. Results of the analysis showed that for the configuration and cases studied, Tollmien-Schlichting (TS) amplification was the dominating disturbance influencing transition. For these cases, incompressible linear stability theory correlated with an n-factor for TS waves of approximately 10 at transition. The n-factor method correlated rather consistently to this value despite a number of non-ideal conditions which indicates the method is useful as a design tool for advanced laminar flow airfoils.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3999 , NAS 1.26:3999
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Computational rotorcraft flowfield research at the U.S. Army AeroStructures Directorate (1990)
    In:  Other Sources
    Details
    Publication Date: 2018-12-01
    Description: Rotorcraft aerodynamic interaction investigations conducted experimentally and computationally by the U.S. Army at the NASA Langley subsonic tunnel are described. Consideration is given to fuselage velocity computation, an interactional aerodynamics method, a rotor/wake/fuselage method, predicted wake geometry, views of tip vortex trajectory, and an inflow velocity study. Calculations were found to predict the periodic geometry of a rotor wake. Details of its interactions with the fuselage surface are not well modeled with inviscid panel methods. For the cases investigated in the present study, the overall effects of the fuselage are relatively small. However, increasing the size of the fuselage or decreasing the fuselage-rotor space will amplify these effects.
    Keywords: AERODYNAMICS
    Format: text
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    NASA TECHNICAL REPORTS
  • 6
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    Navier-Stokes and potential theory solutions for ahelicopter fuselage and comparison with experiment (1994)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A thin-layer Navier-Stokes code and a panel method code are used to predict the flow over a generic helicopter fuselage. The computational results are compared with pressure data at four experimental conditions. Both methods produce results that agree with the experimental pressure data. However, separation patterns and other viscous flow features from the Navier-Stokes code solution are shown that cannot be easily modeled with the panel method.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4566 , L-17325 , NAS 1.15:4566 , ATCOM-TR-94-A-013
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    NASA TECHNICAL REPORTS
  • 7
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    Boundary-layer stability analysis of NLF and LFC experimental data at subsonic and transonic speeds (1987)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: NASA-Lewis has conducted wind tunnel experiments to ascertain the effectiveness of state-of-the-art in natural laminar flow (NLF) and LFC airfoils for subsonic and transonic speeds, such as the NLF(1)-0414F and the SCLFC(1)-0513F. Attention is given to the effects of Tollmien-Schlichting (TS) and/or crossflow linear mechanisms amplifying small disturbances to generate turbulence. It is found that the incompressible TS transitional n-factors were generally in the 9-12 range, in agreement with earlier correlation studies; the TS instability was the dominant instability mode on a swept-planform LFC airfoil over the entire range of test conditions.
    Keywords: AERODYNAMICS
    Type: SAE PAPER 871859
    Format: text
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    NASA TECHNICAL REPORTS
  • 8
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    A multi-element vortex lattice method for calculating the geometry and effects of a helicopter rotor wake in forward flight (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A method is described for the analysis of the unsteady, incompressible potential flow associated with a helicopter rotor and it's wake in forward flight. This method is particularly useful in low advance ratio flight due to the major contribution, in the near field, of the deformed wake. The rotor geometry is prescribed and the unsteady wake geometry is computed from the local flow perturbation velocities. The wake is modeled as a full vortex lattice. The rotor geometry is arbitrary and several rotor blades can be represented. The unsteady airloads on the rotor blades are computed in the presence of the deformed rotor wake by a time-stepping technique. Solution for the load distribution on the blade surfaces is found by prescribing boundary conditions in a reference system which rotates with the blade tips. Transformation tensors are used to describe the contribution of the wake in the inertial system to the rotor in the rotating reference system. The effects of blade cyclic pitch variation are computed using a rotation tensor. The deformation of the wake is computed in the inertial frame. The wake is started impulsively from rest, allowing a natural convection of the wake with time.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0664
    Format: text
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    NASA TECHNICAL REPORTS
  • 9
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    RWF rotor-wake-fuselage code software reference guide (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The RWF (Rotor-Wake-Fuselage) code was developed from first principles to compute the aerodynamics associated with the complex flow field of helicopter configurations. The code is sized for a single, multi-bladed main rotor and any configuration of non-lifting fuselage. The mathematical model for the RWF code is based on the integration of the momentum equations and Green's theorem. The unknowns in the problem are the strengths of prescribed singularity distributions on the boundaries of the flow. For the body (fuselage) a surface of constant strength source panels is used. For the rotor blades and rotor wake a surface of constant strength doublet panels is used. The mean camber line of the rotor airfoil is partitioned into surface panels. The no-flow boundary condition at the panel centroids is modified at each azimuthal step to account for rotor blade cyclic pitch variation. The geometry of the rotor wake is computers at each time step of the solution. The code produces rotor and fuselage surface pressures, as well as the complex geometry of the evolving rotor wake.
    Keywords: AERODYNAMICS
    Type: NASA-TM-104078 , NAS 1.15:104078 , AVSCOM-TR-91-B-008
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    Wind-tunnel evaluation of a 21-percent-scale powered model of a prototype advanced scout helicopter (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An exploratory wind tunnel investigation of a 21 percent scale powered model of a prototype advanced scout helicopter was conducted in the Langley 4 by 7 Meter Tunnel. The investigation was conducted to define the overall aerodynamic characteristics of the Army Helicopter Improvement Program (AHIP), to determine the effects of the rotor on the aerodynamic characteristics and to evaluate the effect of a mast mounted sight on the aircraft stability characteristics. Tests covered a range of thrust coefficients, advance ratios, angles of attack and angles of sideslip and were run for both rotor on and rotor off configurations. Results of the investigation showed that the prototype configuration was longitudinally unstable with angle of attack for all configurations tested. The instability was due to unfavorable interference effects between the horizontal tail and the wake shed from the engine pylon and rotor hub, which caused a loss of horizontal tail effectiveness. The addition of the mast mounted sight had little effect on the stability of the model, but it caused an alteration in the rotor lift distribution that resulted in substantial interference drag for the sight.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2420 , L-15895 , NAS 1.60:2420 , AVSCOM-TR-85-B-2 , AD-A157115
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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